The invention concerns roll and yaw attitude control of a satellite stabilised on three axes in an operational orbit.
In this context a satellite is any artificial object in the solar system:
orbiting the Earth or any other planet in the solar system, or
orbiting a satellite of any planet in the solar system, or
in solar orbit, possibly a transfer orbit between two planets.
The attitude of an orbiting satellite is disturbed by various torques, the major causes of which are:
the asymmetry of the solar radiation pressure due to the angle of the pitch axis (Y) of the satellite relative to the Sun (which angle is not equal to 90.degree.), to the differing reflectivity of different parts of the satellite and to any geometrical asymmetry of the satellite,
the terrestrial gravity gradient,
the terrestrial magnetic field, and
the aerodynamics of the environment (in low orbits).
Consequently, a system for controlling the attitude of a satellite in its orbit is essential. Four types of active system have previously been proposed.
Three of these active attitude control systems increase the mass of the satellite:
the use of thrusters primarily intended for station-keeping: this widely used solution requires an additional mass of propellants for controlling the attitude of the satellite (typically 9 kg of propellants for a satellite life of seven years),
the use of the terrestrial magnetic field interacting with magnetic dipoles created on board the satellite by current loops: this solution requires the provision of coils and in some instances ferromagnetic cores,
the use of the solar radiation pressure acting on specific surfaces that can be deployed and/or oriented relative to the satellite body by actuators: this solution increases the mass of the satellite and reduces its reliability as a result of adding the orientable surfaces and their deployment and/or actuation mechanism.
Representative prior art includes:
patent FR-2.513.589: PROCEDE ET DISPOSITIF POUR ACTIONNER L'AXE DE ROULIS D'UN SATELLITE AVEC UNE DIRECTION DESIREE, PA0 patent DE-2.537.577: LAGERREGELUNG FUR SATELLITEN, PA0 patent FR-2.550.757: REGULATION DE POSITION DE SATELLITES, PA0 patent U.S. Pat. No. 3,304,028: ATTITUDE CONTROL FOR SPACECRAFT, PA0 patent FR-2.529.166: PROCEDE DE MAINTIEN EN POSITION D'UN SATELLITE PAR LA NAVIGATION A L'AIDE DE VOILE SOLAIRE ET VEHICULE SPATIAL METTANT EN OEUVRE LE PROCEDE, PA0 patent FR-2.530.046: PROCEDE ET DISPOSITIF DE COMMANDE D'ATTITUDE POUR SATELLITE GEOSYNCHRONE. PA0 ATTITUDE CONTROL BY SOLAR SAILING--A PROMISING EXPERIMENT ON OTS 2--ESA JOURNAL 1979, Vol 3. PA0 ONE YEAR OF SOLAR SAILING WITH OTS--ESA BULLETIN Aug. 31, 1982. PA0 SYSTEME DE CONTROLE D'ATTITUDE D'UN SATELLITE GEOSTATIONNAIRE--Patent FR-2.531.547. PA0 SYSTEM FOR CONTROLLING THE DIRECTION OF THE MOMENTUM VECTOR OF A GEOSYNCHRONOUS SATELLITE--U.S. Pat. No. 4,325,124. PA0 also U.S. Pat. No. 3,945,148: SATELLITE ROTATION BY RADIATION PRESSURE, which proposes the use of the solar pressure on the solar panels, which are appropriately oriented to rotate the satellite in order to stabilise it. PA0 hold it in the stowed position until the satellite reaches its orbital configuration, PA0 deploy it and maintain it in the deployed configuration, PA0 heatshields which are used to limit heat loss from the satellite during phases in which the solar generator is not fully deployed, PA0 surfaces which improve the luminous flux impinging on the photovoltaic elements (shadow uniformisation screens, for example), and PA0 solar sails designed to modify the system's solar command torque capacity. PA0 1. satellite with low or no kinetic moment and any orientation (in short null or no kinetic moment) provided with three reaction wheels, one on each axis; in this case, the mass saving secured by the invention relates to: PA0 2. satellite with fixed kinetic moment on the pitch axis (Y); in this case the mass saving secured by the invention relates to the roll-yaw control propellants, as this control is no longer provided by thrusters, unless the solar control function leads to a loss of electrical power from the solar generator which is not compensated for by the power margins; depending on the satellite, this occurs for periods accounting for up to 10% of the satellite's life; PA0 3. satellite with kinetic moment on the pitch axis (Y), the kinetic moment being orientable in one direction using any of the following devices, for example: PA0 in this case, the mass saving secured by the invention relates to: PA0 4. satellite with kinetic moment along the pitch axis (Y), the kinetic moment being orientable in two directions using either of the following devices, for example: PA0 two reaction wheels each oriented according to a respective degree of freedom of the kinetic moment, or PA0 an orientable kinetic wheel whose axis is mounted on a double pivot or a universal joint; PA0 in this case, the mass saving secured by the invention relates to the propellants for desaturating the component in the roll-yaw plane of the kinetic moment as the latter is desaturated by the solar control function; PA0 5. satellite provided with continuous actuators (for example: magnetocouplers or electrical propulsion in addition to kinetic wheels or reaction wheels as mentioned previously); in this case the saving secured by the invention relates to:
The fourth type of active system, which is the only known system capable of controlling the attitude of a satellite stabilised on its three axes without incurring a mass penalty, entails orienting the surfaces of the solar panels relative to the Sun by using their drive motors, in order to create torques around two axes perpendicular to the pitch axis (Y) as a result of the effect of the solar radiation pressure. This technique uses equipment already provided on the satellite:
the solar panels, used as the surfaces exposed to the solar radiation,
the solar generator drive motors, used as actuators for these surfaces.
Representative prior art includes:
In the following description the term "solar generator" refers to the combination of both solar panels, the term "solar panel" designating the systems that can be oriented by the drive motors, namely:
the photovoltaic elements of the solar generator,
the structure supporting these elements,
the mechanisms associated with this structure which:
all the additional elements which, in the orbital configuration, are fixed to the structure and which have various roles, including (for example):
In some cases a satellite has deployable heat sinks which can also be used as surfaces exposed to the solar radiation.
The previously mentioned U.S. Pat. No. 4,325,124 discloses an application of this principle which uses data from the terrestrial sensor four times each day to determine the depointing of the satellite relative to S and P inertial axes (see FIG. 1). This data is then used in an open-loop control system to manoeuvre one of the solar panels, advancing or retarding it relative to its nominal displacement facing the Sun, in order to create torques adapted to return the satellite towards the desired attitude.
The previously mentioned patent FR-2.531.547 discloses another application of this principle which uses data from the terrestrial sensor twice each day to determine the depointing of the satellite relative to the S and P inertial axes. This data is then used in an open-loop control system to modify the orientation of the solar panels to compensate inertial torques disturbing the satellite.
The major disadvantages of these control systems which impose no mass penalty are as follows:
the control system is designed to control the satellite relative to an inertial frame of reference which makes it incapable of fine control of a satellite subject to disturbing torques in a satellite-oriented frame of reference and/or representing a second harmonic of the orbital period; also, this control system necessarily has the same pointing performance relative to the roll axis (X) and the yaw axis (Z); this improves the pointing performance in yaw (which is generally not necessary) to the detriment of the pointing performance in roll (which is generally much more necessary);
the attitude is measured only once each day, which gives a very long attitude control response time and considerably restricts the control mode capture field and renders it sensitive to nutation phenomena;
the control system does not take account of the significant correlation between the torques generated about the two axes, making some combinations of torque impossible; this limits the control system to compensating only certain favourable combinations of inertial torques;
it is impossible to benefit from attitude control concepts other than that based on the kinetic moment fixed with respect to the pitch axis (Y) or from the capabilities of other actuators on board the satellite.